Jet engine cooling system



April 1968 o. P. PRACHAR 3,377,803

JET ENGINE COOLING SYSTEM Filed Aug. 10. 1960 5 Sheets-Sheet l IN VEN TOR.

April 16, 1968' o. P. PRACHAR i 3,377,803

J'ET ENGINE COOLING SYSTEM Filed Aug. 10. 1960 5 Sheets-Sheet 2 ATTORNEYApril 16, 1968 o. P; PRACHAR JET ENGINE COOLING SYSTEM 5 Sheets-Sheet 3Filed Aug. 10. 1960 ATTORNEY April 16, 1968 o. P. PRACHAR 3,377,803

JET ENGINE COOLING SYSTEM Filed Aug. 10. 1960 5 Sheets-Sheet 4 INVENTOR.

ATTOR/VE) April 1958 o. P. PRACHAR 3,

JET ENGINE COOLING SYSTEM Filed Aug. 10, 1960 5 Sheets-Sheet 5 IN VENTOR.

United States Patent 3,377,803 JET ENGINE COOLING SYSTEM (Itakar I.Prachar, Birmingham, Mich, assignor to General Motors Corporation,Detroit, Mich., a corporation of Belaware Filed Aug. 19, 1960, Ser. No.48,732

6 Claims. (Cl. 60-261) This invention relates to a fuel and coolingsystem for a gas turbine engine or the like, and more specifically tothe cooling of the different hot sections of a turbomachine by the fuelthat is burned in the combustion chambers to provide the gas propulsiveforce to the engine.

In the past, the design of a high speed air breathing type enginenecessarily included additional cooling equipment because of theinability of the metals used to withstand the extremely hightemperatures prevailing in the combustion and other hot sections of theengine. This, however, adds to the engine weight and often entails theuse of bulky equipment projecting into the airstream thereby creatingadditional skin friction, etc., decreasing the eificiency of the engine.

This invention eliminates these disadvantages by providing an enginewherein the hot sections are cooled by a portion of the fuel that isburned in the combustion sections of the engine. More specifically, thisinvention contemplates cooling the main and afterburner section linersof a gas turbine engine by the absorption of the heat therefrom by .acryogenic fuel of high specific heat capacity, the fuel also cooling theradially inner portions of the compressor rotor, the turbine rotorstage, and the annular innerbody exhaust cone structure.

Therefore, it is an object of this invention to provide a fuel andcooling system for a gas turbine engine or the like utilizing acryogenic fuel of high specific heat capacity for not only burning inthe combustion sections but also for cooling the hot sections of anengine.

Other features, objects and advantages will become apparent uponreference to the succeeding detailed description of the invention and tothe drawings illustrating the preferred embodiment thereof, wherein;

FIGURE 1 is a schematic side elevational view with parts broken away andin section of an engine embodying the invention,

FIGURES 2, 3 and 4 are enlarged views of details'of FIGURE 1,

FIGURE 5 is an enlarged cross-section view of a portion of the FIGURE 1construction taken on a plane indicated by and viewed in the directionof the arrows 5':? of FIGURE 1,

FIGURE 6 is an enlarged cross-sectional view taken on a plane indicatedby and viewed in the direction of arrows 6-6 of FIGURE 1,

FIGURES 7 and 8 are enlarged cross-sectional views of portions ofdetails of FIGURE 1 taken on planes indicated by and viewed in thedirections of the arrows 77 and 38, respectively, of FIGURE 1;

FIGURE 9 is an enlarged cross-sectional view taken on a plane indicatedby and viewed in the direction of the arrows 9-9 of FIGURE 5;

FIGURES 10 and 11 are enlarged cross-sectional views taken on planesindicated by and viewed in the direction of the arrows 16-10 and 11-11of FIGURE 1, and

FIGURE 12 is a perspective view of the details of FIGURE 9.

The drawings, particularly FIGURE 1, show an axial flow air breathingtype engine having a compressor sec tion 10, a main combustion section12, a turbine section 14, an afterburner section 16, and an exhaustnozzle 18. In brief, the invention relates to the use of cryogenic orvery low temperature (-400 F., for example) fuels of high specific heatcapacity for cooling the main and after- 3,377,803 Patented Apr. 16,1968 burner combustion section liners, the internal compressor andturbine rotor structure, and the inner exhaust cone structure. The heatof the parts is absorbed by the fuel which is subsequently injected intothe combustion sections and burned.

With respect to the invention, the construction of the individualsections of the engine will first be described,

with the overall operation of the engine and cooling systems beingdescribed later.

The engine includes at its forward or upstream end an inlet 20 definedby an outer annular engine casing portion 22 and an inner annularconical fairing 24 supported from the casing by a number (only oneshown) of circumferentially spaced support struts 26. Casing 22 isconnected at 28 to the annular engine casing portion 30 0f thecompressor section 10.

The compressor section comprises a four stage axial flow typecompressor, each stage including a rotor wheel assembly 32 and a statorvane assembly 34. The rotor assemblies 32 each have a number ofcircumferentially spaced rotor blades 36 secured to a wheel 38, whilethe stator vane assemblies 34 each have a number of circumferentiallyspaced stator vanes 40- secured to and through the engine casing portion30 by a supporting ring 42 hatshaped in cross-section. The rotor bladesof all the stages have annular outer shrouds 43 sealingly cooperatingwith stationary seal portions 44 secured to the casing 30. The statorvanes of the first three stages have annular inner shrouds 4-5 havingstationary seal portions 46 cooperating with rotating seal portions 47secured to stiffening drums 48. The drums are positioned between andsecured to the rotor wheels to interconnect the four stages of thecompressor rotor for simultaneous rotation. The rotor stages arerotatably supported within the engine casing by the mounting of thefirst stage rotor wheel 49 on the inner race of a bearing means (notshown), the outer race being secured to an extension of the supportingstruts 26.

Extending downstream of the compressor section is a diverging airdiffusion passage 50' defined by an outer engine casing portion 51 andan inner annular wall 52. Wall 52 is secured to an inner annular shroud53 fixed to the circumferentially spaced guide vanes 40 secured to theengine casing in the manner shown.

The main combustion section comprises in general a single annularchamber bounded on the outside by the annular engine casing and on theinside by an annular wall. Each of the chamber walls is of a hollowdouble wall construction so that cryogenic fuel can be pumped throughthe hollow annulus from one end of the chamber to the other to cool thecombustion chamber lining. The fuel is then injected into the combustionchamber proper and burned along with the main supply of fuel in thechamber.

As seen best in FIGURE 4, both the outer and inner diffuser sectionwalls 51 and 52 are extended further downstream to become a part of theouter and inner walls 54 and 55', respectively, of the combustionchamber 12. Both walls are of a sandwich-type construction (FIGS. 7 and8), each consisting of two parallel annular walls radially separatedfrom each other by an annular circumferentially corrugated sheet metalstrip.

The construction of the outer chamber wall 54 consists of the twoparallel walls 56 and 58 which extend for the length of the chamber andare joined to each other at opposite ends. The corrugated strip 60 formshollow separated cooling passages 62 between each of the walls to whichit is connected and opposite sides of the strip. It extends for almostthe entire length of the combustion chamber but is terminated short atboth ends to form annular manifolds 64 and '65, respectively, betweenthe ends of the strip, and the joined wall edges.

As will be described in more detail later, fuel is fed into manifold 66,passes through the cooling passage 62 and is again collected by themanifold 64 to be fed into the combustion chamber through a number ofspray bars 68. The fuel thus cools the walls 56 and 58 which itcontacts, and the manifold 64 acts as both a collector and distributor.Manifold 64 also serves another purpose, and that is to act as aconveying manifold for the main flow of fuel directly from a main fuelsupply line (not shown) into the main combustion chamber fuel spray bars70.

To perform both of the above functions and yet maintain both of the fuelflows separated, the manifold 64 is divided into two parts 71 and 72 bya partition or wall 73 extending completely around the engine, as bestseen in FIGURE 12. Manifold 71 is fed fuel directly from another annularfuel manifold 74 through a number of holes 76 in the outer casing wall56, and feeds the fuel directly therefrom through holes 78 in the outerliner wall 58 into the open end of the circumferentially spaced mainfuel spray bars 70. Spray bars 70 are each secured to wall 58 directlybeneath the openings, and may be tubular as shown with a number of fuelorifices St) in the casing equally spaced along the length of the bar.Since they are the main spray bars, they extend substantially across theentire width of the combustion chamber to evenly distribute the fuelspray into the chamber. Fuel manifold 74 is connected at one edge towall 56 and at the other edge to the diffuser casing 51. It is fed fuelfrom the main fuel supply line (not shown) by a connecting line 82.

The other part 72 of the manifold 64 formed by the partition 73cooperates with its own fuel spray bars 68 for delivering the coolingfuel from passages 62 into the combustion chamber. Spray bars 68 arecircumferentially spaced from each other and the main spray bars 79, butare in the same radial plane to better mix the fuels. Spray bars 68 arealso directly beneath manifold 71. Therefore, in order for the fuel topass from the manifold 72 into spray bars 68, it must pass through aportion of manifold 71. To accomplish this without mixing the fuelflows, a number of somewhat tunnel-like constructions 83 are provided inmanifold 71, whereby in each instance, the fuel in manifold 72 passesthrough a hole 84 in the wall 73 and into a trough-shaped conduit orheader 86. The header 86 directs the fuel into spray bar 68 while at thesame time isolating the fuel from the fuel in manifold 71. Header 36 isclosed at one end 88 as seen in FIGS. 9 and 12, and is joined at itsopposite open end to the partition 73. The walls of the header aresecured to wall 58 thereby completing the three-sided hollow conduit.Holes 90 are provided in wall 58 directly beneath each header, and thefuel is fed through the holes into the open end of the fuel spray bar 68secured to the wall. Spray bars 68 being secondary spray bars extendonly partially into the combustion chamber. As best seen in FIG. 5, thespray bars 68 are of a construction similar to that of the main spraybars 70 and, as stated before, are spaced between. The number of headers86, the spacing thereof, and the number of holes in wall 73 correspond,of course, to the number of spray bars 68. Thus, the cooling fuel inpassages 62 passes into spray bars 68 through the partition 73 andheaders 86 without mixing with the fuel in manifold 71.

A brief fuel flow description will sum up this phase of the operation ofthe engine. The entire fuel to be used in the engine initially flowsfrom a source (not shown) of very low temperature (400 R, for example)cryogenic liquid fuel under pressure having a high specific heatcapacity. The fuel may or may not be fed into a heat exchanger (notshown), which if it is changes the liquid fuel to its vapor form beforepassing it through the engine through the main supply line. if it is notpassed through a heat exchanger, the liquid fuel will vaporize almostimmediately upon absorption of heat from the system since thevolatilization point is very low. The branch line 82 therefore feeds themain supply of main combustion chamber fuel into manifold 74 and intospray bars 7% through manifold 71. Simultaneously fuel is fed from line32 into a smaller flexible branch line 92 connecting at its downstreamend with the manifold 66 (FIG. 3) and the inlet to passages 62. The fuelthen passes through the cooling passages 62, into manifold 72, throughholes 84 in wall 73 and into spray bars 68 through headers 86. The outercombustion liner wall 58, and casing wall 56 if it becomes hot for somereason, thus gives up the heat absorbed by it during the combustionprocess to the extremely cold fuel vapor, and the liner is maintainedcool.

The cooling of the inner combustion chamber wall 55 is accomplished in amanner similar to that of cooling the outer wall 54, just described. Asstated previously, the inner wall is also of a sandwich-typeconstruction consisting of the two parallel walls 94 and 96 radiallyseparated by a circumferentially corrugated annular sheet metal strip 98to form cooling passages 100 between the strip and walls. The wallsextend for the length of the chamber and are joined to each other attheir upstream ends. The corrugated strip 93 extends for substantiallythe entire length of the chamber but is terminated short of the ends toprovide annular fuel manifolds 192 and 104 at opposite ends of thechamber.

In brief, cooling fuel is fed through a conduit 106 and a connectingconduit 108 to the manifold 10d and the downstream inlet ends of thepassages 109. The fuel then passes upstream cooling the inner linearcombustion liner wall and is collected by the manifold 1G2. Manifold 192directs the fuel through a number of holes 116 in liner 94 into the openends of a number of tertiary fuel spray bars 112 to be distributed intothe combustion chamber. Spray bars 112 are circumferentially spaced fromeach other and staggered with respect to the other spray bars as bestseen in FIG. 5, and are mounted in the same radial plane as the others.Each of the spray bars may be of a construction similar to spray bars70, i.e., a tubular closed end piece with spaced orifices and having itsopen end secured to liner wall 94 over the hole 110. Each of the spraybars 112 extends approximately to the chamber centerline.

At the downstream ends of the combustion chamber, the walls 94- and 96are both bent radially inwardly in several different directions as seenin FIG. 3. The walls pass around the inner shroud 113 of the turbineinlet nozzle 114 and substantially parallel the curvature of the turbinerotor wheel 115 and wheel-to-blade connection 116 to aid in cooling theface of the turbine rotor in a manner to be described later. Theradially innermost inlet end 118 of the manifold 104 opens into and isfed fuel from the fuel supply conduits 106 and 16% previously mentioned.Fuel line 166 extends through and spans the diffusion passage 50 and isconnected directly to the fuel manifold 74 through an opening 120 toreceive the fuel from the supply line 82.

As thus far described therefore, the cryogenic fuel flows in threepaths; first, the main fuel supply in manifold 74 flows directly intothe primary fuel spray bars 76 through holes 76 in wall 56, manifold 71and holes 78 in wall 58; secondly, from the manifold 74 through flexiblepipe 92 into manifold 66, into and through the passages 62 cooling theliner 58, into manifold 72 and out through headers 86 into spray bars68; and thirdly, through supply lines 106 and 108 to manifold 10% andthe passages 100 to cool the combustion chamber inner liner 94, and theninto the spray bars 112.

As best seen in FIG. 3, the turbine section 14 has an outer wallconsisting of an annular casing portion 122 suitably connected at itsopposite edges to the combustion and afterburner section casings. Theinner wall is radially and axially defined by the inner shroud 113 ofthe turbine inlet guide vanes, the rotor blade platforms 123, and aninner shroud 124. Shroud 124 is connected to an outer annular shroud 126by a number of circumferentially spaced streamlined fairings 128. Eachof the fairings 128 encloses and is secured to one of a number of (onlyone shown) circumferentially separated turbine bearing support struts130 secured to the outer casing 122. Shroud 124 has an annular partition13 2 depending therefrom having an enlarged annulus 133 on its inneredge fixed to the outer race 134 of an annnular bearing 136. The innerrace 138 of the bearing rotatably supports and is secured to theextended hub 140 of the wheel 115 of a single stage axial flow typeturbine rotor 142. A number of circumferentially spaced rotor blades 144are secured to the rotor wheel 115 by any suitable means and cooperatewith a number of circumferentially spaced stator inlet guide vanes 146.Vanes 146 are secured to the casing 122 at one end as shown and at theother end are secured to and support the downstream end of thecombustion liner w-all 94 and inlet end of fuel passages 100.

As stated previously, the different compressor rotor stages are drivenby the turbine through a suitable connection, this being the annularconnecting drive drums 48 between compressor rotor wheels 38 (FIG. 4)and an annular drum 148. Drum 148 is secured at one edge to the laststage compressor wheel 150 and extends to and is secured to the turbinerotor wheel 115 at its other edge by circumferentially spaced bolts 152.The bolts also support and secure the rotatable portion 154 of alabyrinth seal 156 to rotor wheel 115. A conduit 157 maintains the spacebetween the bearing 136 and the seal vented of any gas leaking throughthe seal, the vent leading to an opening (not shown) to vent the gasoverboard of the engine.

The face of the turbine rotor wheel 115, and blade attachment portion116 are cooled by the flow of fuel through manifold 104. This isaccomplished by funneling or directing higher pressure and coolercompressor discharge air to the turbine rotor from the space 158 betweenthe last stage rotor blades 36 and stator vanes 40. The directing meansis defined by an annular gas chamber 159 between the inner diffuser andmain combustion chamber walls 52 and 55 and the drive drum 148. The airflows in the direction indicated by the arrows 168 and passes over thefuel supply conduit 1% and into the space 162 between the rotor wheel115, connection 116 and the extension of the inner liner wall 94 andmanifold 104. The cooling fuel in manifold 104 therefore absorbs theheat from wall 94 which in turn has absorbed the heat from the airrushing past the face of the hot turbine rotor. A number of swirlproducing baffles or fins 164 are attached to the liner 94 to helpcirculate the air and therefore constantly cool the air more eifectivelyby the continuous contact with the cooled walls of the fuel manifold104.

The stiffener drums 48 and 148 not only constitute turbine driventransmitting members to drive the compressor, but also double as theouter wall of an annular cooling annulus or chamber 166 to cool theinterior parts of the compressor. The inner wall of the annulus orchamber is defined by a tubular dual fuel tube or conduit 168 mounted onthe compressor axis. Conduit 168 sealingly engages and is secured to thefirst stage compressor rotor wheel 49 at its upstream end as shown toclose one end of the chamber, and has a sealing engagement at its otherend with the turbine rotor wheel hub 140 in a manner to be described.

In general, the conduit comprises concentrically and telescopicallymounted interconnected tubes, one to carry the cooling fuel in to bedischarged into chamber 166, and one to carry it out. The tubes rotateat turbine rotor speeds and the inlet tube is provided with openings topermit the fuel to be discharged into the chamber under the greatcentrifugal force imposed on the fuel. Once it enters the chamber, thefuel is immediately warmed by contact with the Warm faces of thecompressor rotor and turbine rotor wheels, causing convection currentsto be set up thereby distributing the cool fuel to all parts of thechamber. Since the cool fuel displaces the hotter lighter fuel, the highcentrifugal force of the fuel drives the warmed air to a toroidal-likefashion to all corners of the chamber and out through outlets in theoutgoing tube. The hotter fuel is then discharged into a large coolingconduit to be described.

More specifically, the conduit 168 comprises a central tube 170extending longitudinally through the rotor wheel hubs from itsconnection at 172 to the first stage rotor wheel 49 to a positionadjacent the turbine rotor bearing 136. A sleeve tube 174 telescopicallysurrounds the tube 179 for its entire length and is radially separatedfrom and secured to the tube by five axially separated sets ofstiffeners 176 (FIG. 6). Each stiffener set comprises six hollowreinforcing members 177 spaced from each other and secured at oppositeends to and extending radially between the tubes 176 and 174. The hollowpassage 178 defined by tube 170 constitutes a fuel inlet line, and thespace 179 defined by the annulus between tubes 170 and 174 constitutes afuel outlet line.

To discharge the cooling fuel from passage 178 into chamber 166, anumber of open end hollow plugs or spigots 18G extend radially betweenthe tubes 170 and 174 within the reinforcing members 177. The spigotsare each secured at one end to tube 178 in relieved openings 181, and atthe other end to the wall of sleeve 174 in an opening therein. Each ofthe spigots is located adjacent the faces of the rotor wheels, and theouter ends are partially covered by an annular fuel directing bafile orsplatter shield 182 secured to tube 174. The splatter shields preventthe discharged cooling fuel from striking the rotor wheel hubs directlyto prevent cracking of the hubs.

Sleeve tube 174 at its upstream end is joined to tube 170 closing theend of the passage 179. The connection of tube 178 to the compressorrotor wheel 49 also closes this end of tube 170. Therefore, a number ofholes 183- and 184 are provided in this end of the tubes 178 and 174,respectively, to permit the cool fuel to be discharged into the chamberat this end on the one hand, and to permit the exit of the hotter fuelfrom the chamber on the other hand. The wall of tube 174 also has anumber of fuel exit openings 186 along its length to permit the entryinto the space 179 of the hotter fuel forced from chamber 166 as aresult of the convection currents set up by the centrifuged cooler fuelin the chamber as previously described.

Tube 170 at its downstream end is relieved as shown at 187 and formed toprovide one portion 188 of a labyrinth seal 190. Tube 174 at this end isfixedly inserted into a hollow flanged accessory drive shaft 192, whichis secured to the turbine rotor wheel hub at one edge and has therotating portion 19 of another labyrinth seal 1% formed on its otheredge. Both rotating portions of the seals cooperate with stationary sealportions 198 and 208 formed on a stationary annular support 202. Thesupport is secured to partition 132 at one end, and at the other end tothe end of an expandable fuel tube 204. The

seals therefore maintain the conduit 168 divided into the 7 separatefuel supply and return passages 178 and 179. The portion between theseals is apertured at 206 for discharging the warmed fuel from passage179. The hollow space between seal 198 and a bearing seal 207 may befilled with an inert gas through conduit 208, the gas being at apressure to maintain a constant leakage in a direction through seal 196-to prevent contamination of the bearing with the fuel in conduit 179.

Fuel passage 178 is connected at its inlet end to expandable tube 294which in turn is connected at its opposite end to the bottom or one endof an enlarged fuel distributing pipe 214 secured within one of thehollow support struts 130. Pipe 214 is adapted to be connected at itsother end through an opening in the engine casing '7? 122 to the samesource (not shown) of cryogenic fuel as that supplying the fuel for themain combustion chamber, i.e., a very low temperature fuel having a highspecific heat capacity. The fuel in this pipe 214 serves as the fuel forafterburning as well as for cooling, as will be described, and thesource of this afterburner fuel may be entirely separate from the maincombustion chamber fuel source, if desired.

With the construction as thus far described, the chamber 166 issubstantially sealed. The cooling of this por tion of the engine istherefore as follows. Under normal running conditions of the engine,admission of the afterburner fuel into pipe 211 i causes fuel to flowinto line 2% and into the inlet of passage 37% to tic-w axially upstreamunder pressure. Since the tube is rotating at turbine rotor speed, thefuel will be centrifuged out into chamber 166 through the spigots 13%with great force, the fuel being deflected by the deflector 183 toprevent a stream of the extremely cold fuel from directly hitting therotor hubs. The centrifuged fuel flows radially outwardly as indicatedby arrows 216 along the f :es of each of the compressor and turbinerotor wheels 35 absorbing the heat thereof. As indicated by the arrows218, the fuel then flows and is circulated to all parts of the chamberby means of convection currents set up by the displacement of the warmedfuel by the cooler centrifuged fuel. The warmed fuel is then forced outof the chamber by the cooler fuel and into the return annulus 179through the outlets 186 in tube 174. The fuel then flows through line179 and is finally discharged out through openings 206. Thus, the facesof the rotor wheels and the drums 4% and 148 and other parts of thechamber are cooled by the fuel absorbing the heat from the parts.

Returning now to the construction of the remaining sections of theengine, the afterburner com ustion chamber extends rearwardly from theturbine section and is delined by an outer annular engine casing 22?-and an inner annular wall 224 formed as a part of the periphery of theinner exhaust cone 226. Both the outer casing and inner wall are of asandwich type construction similar to that of the main combustionsection to provide cooling passages through which cooling fuel is forcedto cool the afterburner combustion liner and the inner exhaust cone. Theouter casing 222 consists of two parallel walls 223 and 23f) radiallyseparated from each other by a circumferentially corrugated annularsheet metal strip 232 to form cooling passages 234 between the stripsand Walls. The walls extend for the length of the section and are joinedto each other at opposite ends. The corrugated strip 232 extends foralmost the entire length of the section but is terminated short of theends to provide annular manifolds 236 and 238 between the ends and thestrip.

Cooling fuel is fed from pipe 214 through the expandable branch line 24%into an annular manifold 1242 surrounding manifold 238 and secured towall 228. The fuel then flows into manifold 238 through a number ofcommunicating holes 243 in wall 2%, and flows upstream through passages234 into manifold 236, cooling the liner wall 239. Manifold 236 thendirects the fuel through holes 244 in wall 230 into the open ends of anumber of secondary afterburner spray bars 245 secured to wall 23% andlocated at the inlet to the afterburner combustion chamber. The spraybars are of a construction similar to spray bars iii, arecircumfcrentially spaced from each other, and extend approximately tothe center line of the afterburner section.

As stated previously, the inner afterburner chamber wall 224 is formedas an integral part of the outer periphery of the exhaust nozzle innercone as seen in FIG. 1. Except for the particular converging anddiverging shape, the outer confines of the wall are, however, of asandwich type construction similar to that of the outer wall The innerwall is constructed in such a manner that the main supply of theafterburner combustion chamber fuel flows from pipe 214 downstreamthrough a passage in the center of the exhaust cone to a point close tothe of the cone, where the passage then is turned approximately 180 toflow upstream along the periphery of the cone and afterburner innercombustion liner to cool the same. The fuel is then discharged into theafterburner combustion chamber proper through primary fuel spray bars.

More specifically, inner wall 224 consists of two parallel walls 246 and24S radially separated by an annular corrugated strip 7.50 secured toboth to provide hollow cooling passages 25?. between the strip andwalls. The walls are joined to each other and to the turbine sectioninner shroud 124 at the discharge end of the fuel passages, and thecorrugated strip is terminated short of this end connection to provide amanifold 253 therebetween. The manifold directs fuel from the passagesthrough a number of holes 254 in wall 246 into the open ends of a numberof primary afterburner spray bars 2% secured to the wall. The spray barsare of the same construction as spray bars 7%), and extend substantiallyacross the entire width of the combustion chamber. They are in the sameradial plane as secondary spray bars 244 but are staggeredcircumferentially therefrom and each other in the same manner as spraybars 68, 7b and 112 to provide the most efficient and even distributionof fuel into the combustion chamber.

Both walls 246 and 24S and the corrugated strip 250 extend downstream toa point on the cone as shown at 257, the corrugated strip terminating atthis point. The outer wall 2-;6 extends further downstream to form theapex of the cone. Wall 248 at this point turns inwardly together with anannular guide member 258 form a radial cooling manifold or passage 259.Wall 248 is again then turned 90 as shown while guide member 258diverges to form between the two a large fuel passage 260 extendingupstream as shown. The fuel inlet end of passage 26% is again enlargedby flaring the end portion of wall 248 and securing it to the connectionbetween the turbine inner shroud 12-4- and the end connection of walls246 and 248. The inlet end of passage 260 thus defines together with anannular stiffener 262 secured to pipe 214, the annular partition 132,and the support 262 a fuel cooling chamber 264. The main supply of fuelfor distribution through the primary afterburner fuel spray bars 256,while also is the cooling fuel, is fed into chamber 264 through a numberof apertures 266 in the bottom or radially inner end of pipe 214-. Thefuel discharged from the annulus 179 of conduit 168 also is fed intothis chamber. Thus, the cryogenic fuel flows along passage 2% andoutwardly into passages 252 to flow upstream to cool the innerafterburner liner as well as the exhaust cone. The fuel then flows intomanifold 253 and out into the combustion chamber through spray bars 256.It is to be noted that the outer exhaust nozzle portion includes anannular shroud 263 secured at its forward end as shown to the edge ofthe afterburner outer casing and manifold. This outer exhaust nozzleportion also may include relatively movable elements 270 (only partiallyshown) for varying the area of the nozzle. Other details are not givenbecause they are known, form no part of the present invention, and areimmaterial to an understanding of the invention. t

It is also to be noted that the particular shapes and configurations ofthe main and afterburner combustion sections and exhaust nozzle andother sections are of a design according to known practices to producethe most effiient operation of the engine.

Having described the details of the invention, a brief description ofthe overall operation of the engine and the cooling system therefor isas follows.

Under normal running conditions of the engine, air taken in through theinlet Ztl passes into and through the co npressor from which it isdischarged with a higher velocity, temperature, and pressure into thediffuser section wherein the velocity is reduced and kinetic energyconverted to static pressure. From the diffuser section, the

air is discharged into the main combustion chamber where it is mixedwith cryogenic fuel fed into the chamber from the fuel spray bars 70, 68and 112. Ignition of the mixture by any suitable means (not shown)causes the turbine 142 to be driven by the products of combustion, themotive fluid discharged through the turbine passing into theafter-burner combustion chamber. There it is combined with the cryogenicfuel vapor sprayed into the chamber from the fuel spray bars 245 and 256and ignited by suitable means (not shown) to produce the additionalthrust desired, the exhaust nozzle further increasing the velocity,etc., of the gas.

- Further details as to the general operation of the engine beyond thosedescribed are believed to be unnecessary since they are known and areimmaterial to an understanding of the invention.

As to the operation of the cooling system, the main combustion chambercryogenic fuel at 400 F., for example, is fed for burning purposesinitially from the main supply line (not shown) into the main fuel spraybars 70 through the fuel inlet 82, manifold 74, and manifold 71.Simultaneously, the cryogenic fuel is fed for cooling purposes frommanifold 74 into branch line 92, manifold 66 and into the inlet ofcooling passages 62 to flow upstream along the length of the maincombustion chamber outer liner. The liner is maintained cool by the fuelabsorbing the heat therefrom, and the fuel is discharged from thesecondary fuel spray bars 68 into the combustion chamber.

Main combustion chamber fuel is also fed from manifold 74 into conduits106 and 108 to manifold 104 and the inlet end of passages 100 to coolthe inner combustion liner 55, the fuel being subsequently dischargedinto the combustion chamber through spray bars 112. The cooling fuel inmanifold 104 also maintains the compressor discharge air flowing pastthe turbine rotor face in chamber 162 cooled by contact of the air withthe manifold.

Thus, the main combustion chamber fuel not only supplies the combustionchamber with a main source of fuel but also cools the inner and outerliners of the chamber as well as the face of the turbine rotor wheel.

A portion of the after-burner cryogenic fuel in pipe 214 passes intoconduits 204 and passage 178 of tube 170 from which it is centrifugedout into chamber 166 to cool the rotor wheel faces and the internalcompressor and turbine parts enclosed thereby. The fuel then is forcedinto return annulus 1'79 and out into chamber 264 to maintain thebearings, etc., cooled. The cooling of the drive drum 148 also aids incooling the compressor discharge air passing along the outer sidethereof.

The fuel discharged from the annulus 179 is combined with the mainsupply of afterburner fuel fed into chamber 264 from the apertures 266in the bottom of pipe 214 to the inlet of fuel passages 252. The fuelthen flows upstream cooling the inner exhaust cone casing and innerafterburner combustion liner and is subsequently discharged into theafterburner chamber through the main afterburner fuel spray bars 256.Simultaneously, the afterburner fuel is fed from pipe 214 into thebranch line 240 and therefrom into the inlet end of the afterburnerouter cooling passages 234 to flow along the same and cool the outercombustion liner 230. The fuel is then discharged into the afterburnerchamber through spray bars 250,

The cryogenic fuel therefore is not only the main supply of fuel forburning in the main and afterburner combustion chambers, but also is acoolant to cool the inner and outer liners of the chambers, as well asthe internal parts of the compressor and turbine rotor wheels and theinner exhaust nozzle. Thus, all hot sections of the engine are cooled bythe passage of extremely cool fuel therethrough having a high specificheat capacity to absorb the heat therefrom.

From the foregoing, therefore, it will be seen that this inventionprovides an engine construction wherein the fuel serves a dual purposeof being both a fuel and a coolant 10 for cooling the walls of thecombustion chambers and other hot sections of the engine, thuspermitting the design of higher Mach No. engines.

While the invention has been illustrated in its preferred embodiment ina gas turbine engine of the air breathing type, it will be obvious tothose skilled in the art to which this invention pertains that manymodifications may be made thereto without departing from the scope ofthe invention.

What is claimed is:

1. A fuel and cooling system for a turbomachine of the axial flow typehaving an outer annular longitudinally extending engine casing and innerlongitudinally extending annular wall means cooperating therewith totogether define axially aligned compressor, main combustion, turbine,and afterburner combustion sections therebetween for the flow of gastherethrough, a plurality of fuel injectors adapted to contain fuelsecured to said casing and wall means and extending into each of themain and afterburner combustion sections for the injection of fueltherein, separate sources of cryogenic fuel each having a high specificheat capacity for said main and afterburner sections, means connectingsaid sources and said main and afterburner combustion section fuelinjectors, portions of said engine casing and said wall means havinghollow interiors adapted to contain fuel under pressure, and meansconnecting said sources of fuel and the hollow interiors of said casingand wall means for the flow of cryogenic fuel under pressuretherethrough to cool the enclosure defining surfaces of said main andafterburner combustion sections, annular means enclosing the interior ofsaid compressor section, and means connecting said afterburner sectionsource of fuel to said last mentioned means for the flow of cryogenicfuel therethrough to cool said interior of said compressor section bythe absorption of heat by said fuel.

2. A fuel and cooling system as in claim 1 wherein said compressorsection has a compressor therein having a rotatable shaft securedthereto, and said means connecting said afterburner source of fuel andsaid compressor interior comprises fuel distributing means including therotating shaft of said compressor.

3. A fuel and cooling system as in claim 2 wherein said shaft is hollowfor containing said fuel and has a plurality of fuel exits therein incommunication with the interior of said compressor, said fuel beingconnected to said hollow shaft at one end thereof, the rotation of saidshaft centrifuging said fuel out through said exits into the interior ofsaid compressor and against said annular means to cool the same by theabsorption of heat by said cryogenic fuel, and means to connect saidfuel from the interior of said compressor to said hollow interiors ofsaid afterburner combustion section inner wall means for conveyance tosaid afterburner fuel injectors.

4. A fuel and cooling system as in claim 1 wherein conduit means areprovided connecting a portion of the gas discharged from said compressorto said turbine section to cool a portion of a turbine containedtherein, said conduit means and said inner main combustion section wallmeans containing cryogenic fuel and extending into adjacency with eachother for absorbing heat from said turbine cooling gas by the saidcryogenic fuel to further cool said turbine portion cooling gas.

5. A combustion chamber cooling and fuel system comprising annularhollow fuel carrying outer and inner walls extending longitudinally andbeing spaced radially from each other to provide a combustion chambertherebetween, a plurality of hollow fuel injectors spaced around andprojecting into said chamber, a source of cryogenic fuel of highspecific heat capacity under pressure, and means to connect the fuelfrom said source to said injectors and to the interior of saidcombustion chamber Walls to cool said walls by the absorption of theirheat by said fuel, said means comprising a first annular fuel manifoldconnected to said source of fuel and surrounding one end of said outerwall, means connecting a portion of said manifold and the interior ofsaid inner wall at one end, means connecting the other end of said innerwall and some of said fuel injectors for the flow of fuel from saidmanifold through said inner wall to said injectors to cool said innerwall, the end of said outer wall within said fuel manifold being closedby means therein defining a closed second annular fuel manifold, saidsecond manifold having apertures therein some of which connect the fuelfrom said first manifold directly into and through the second manifoldand into the open ends of others of said fuel injectors secured to saidsecond manifcld, and fuel conducting means in said second manifoldcommunicating at one end through others of said apertures in said secondmanifold with the fuel in the hollow outer wall and at their other endsthrough still others of said apertures with the open ends of furtherones of said fuel injectors secured to said second manifold, and meansconnecting the fuel in said first manifold to the opposite end of saidouter wall for flow therethrough and into said further ones of saidinjectors, the fuel fiow through said walls cooling the walls andvaporizing the fuel.

6. A fuel cooling and supply system for a turbomachine having acompressor, a main combustion section, a turbine, an afterburnersection, and an exhaust nozzle, said compressor having hollow meanstherein enclosing the interior of said compressor forming a chambertherein and being drivenly connected to said turbine, and fuelconducting means in said chamber connected to said turbine, saidconducting means comprising a plurality of nested cylindrical hollowmembers each having a plurality of openings therein and being radiallyseparated to form separated conduits, a source of cryogenic afterburnerfuel under pressure having a high specific heat capacity, meansconnecting said source to the end of the inner of said conduits, meansprojecting from the openings in said inner member through the remainingof said plurality of members to open into said chamber for centrifugingthe cryogenic fuel into said chamber against said hollow means uponrotation of said members to cool the hollow means by the absorption ofits heat by said fuel, said cryogenic fuel flowing out of said chamberthrough the openings in the remaining of said plurality of members, aplurality of afterburner fuel injectors, said exhaust nozzle being ofthe plug type, said plu having hollow wall means defining the peripherythereof connected at one end to said fuel injectors, and meansconnecting the fuel flowing out said fuel conducting means to theopposite end of said plug wall means to cool the plug wall means by theflow of fuel therethrough to the fuel injectors.

References fired CARLTON R. CROYLE, Primary Examiner.

SAMUEL FEINBERG, Examiner.

D. HART, Assistant Examiner.

UNITED STATES PATENT OFFICE CERTIFICATE OF CORRECTION Patent No.3,377,803 April 16, 1968 Otakar P. Prachar It is certified that errorappears in the above identified patent and that said Letters Patent arehereby corrected as shown below:

Column 6, line 5, "to", first occurrence, should read in Lolumn 7, line53, "the" should read an Column 8, line 32, "inwardly. together" shouldread inwardly to together line 45,

'while" should read which Signed and sealed this 3rd day of March 1970.

SEAL) kttest:

1.1m; M. Fletcher, 1r. WILLIAM E, SCHUYLER, JR.

Ittesting Officer Commissioner of Patents

1. A FUEL AND COOLING SYSTEM FOR A TURBOMACHINE OF THE AXIAL FLOW TYPEHAVING AN OUTER ANNULAR LONGITUDINALLY EXTENDING ENGINE CASING AND INNERLONGITUDINALLY EXTENDING ANNULAR WALL MEANS COOPERATING THEREWITH TOTOGETHER DEFINE AXIALLY ALIGNED COMPRESSOR, MAIN COMBUSTION, TURBINE,AND AFTERBURNER COMBUSTION SECTIONS THEREBETWEEN FOR THE FLOW OF GASTHERETHROUGH, A PLURALITY OF FUEL INJECTORS ADAPTED TO CONTAIN FUELSECURED TO SAID CASING AND WALL MEANS AND EXTENDING INTO EACH OF THEMAIN AND AFTERBURNER COMBUSTION SECTIONS FOR THE INJECTION OF FUELTHEREIN, SEPARATE SOURCES OF CRYOGENIC FUEL EACH HAVING A HIGH SPECIFICHEAT CAPACITY FOR SAID MAIN AND AFTERBURNER SECTIONS, MEANS CONNECTINGSAID SOURCES AND SAID MAIN AND AFTERBURNER COMBUSTION SECTION FUELINJECTORS, PORTIONS OF SAID ENGINE CASING AND SAID WALL MEANS HAVINGHOLLOW INTERIORS ADAPTED TO CONTAIN FUEL UNDER PRESSURE, AND MEANSCONNECTING SAID SOURCES OF FUEL AND THE HOLLOW INTERIORS OF SAID CASINGAND WALL MEANS FOR THE FLOW OF CRYOGENIC FUEL UNDER PRESSURETHERETHROUGH TO COOL THE ENCLOSURE DEFINING SURFACES OF SAID MAIN ANDAFTERBURNER COMBUSTION SECTIONS, ANNULAR MEANS ENCLOSING THE INTERIOR OFSAID COMPRESSOR SECTION, AND MEANS CONNECTING SAID AFTERBURNER SECTIONSOURCE OF FUEL TO SAID LAST MENTIONED MEANS FOR THE FLOW OF CRYOGENICFUEL THERETHROUGH TO COOL SAID INTERIOR OF SAID COMPRESSOR SECTION BYTHE ABSORPTION OF HEAT BY SAID FUEL.